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The prototype of any category is the member or set of members of a category that best represents the category as a whole. Not everything fits perfectly in a category. Categories are defined by an intersection of properties that make up their members. The language is set of different categories and we should attend to categories in language teaching. Gender and mother language are among these components that should be considered in the teaching and learning process. Research about prototype is necessity because no basic prototype study, and the impact of gender and mother language on the formation of this concept have not been made among farsi lernears. As a result, one group of forty speakers who were learning Farsi selected. Each group included 20 men and 20 women. Men from Hojjatieh school and women from the Bent-Alhoda center were selected, both under the supervision of the Almustafa international university. In the category of domestic animals, 41.17% of men chose cows, while the same percentage of women chose sheep, horses, and dogs. In the field of sports, in addition to the first choice, the second and third choices of women and men were completely different, men have chosen shooting and cricket, and women have chosen volleyball, football and walking. Also, in the category of vehicles, although both sexes chose airplane as their first choice, their next choices were different, men chose train and car while women chose bus, car and bicycle.
 

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In this paper, a new method based on Pareto Simplex Search (PSS) has been employed to find the communications satellite optimal transfer trajectory from geosynchronous transfer orbit (GTO) to geosynchronous orbit (GEO) as the destination operational zone. We solve this problem through using a mathematically efficient algorithm and in all transferring phases, considering transfer orbits to intermediate orbits; it is supposed that the orbital maneuver control unit has continuous performance characteristics. The system governing equations representing the two body system of a spacecraft and primary gravitational source and thrust force applying to spacecraft, are defined in a nonlinear form. In this systematic approach we incorporate system dynamics as the problem constraints, and both minimum-time and fuel consumption using constant acceleration simultaneously as the problem strategy to find the optimal transfer trajectory between two orbits. Set of optimal trajectories are plotted in Pareto Front and transfer trajectory can be selected from these points.

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Layout design of a complex system such as a satellite or an airplane is a difficult problem to solve. This problem is known as an NP-complete problem in terms of computational complexity. The main difficulties encountered in the layout design problem are formulation of the problem in mathematics and the solution strategy and practical approaches in engineering. This paper reviews the research work of the authors to automatic layout design process in satellite in last 10 years and the methods of layout design of satellite divides to three main categories that consist of practical method, human-computer interaction and three dimensional layout design and then studies methods of the current state of the art of 3D layout design

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This paper presents a new method to design optimal thrusters’ configuration for geostationary satellite in order to reduce the fuel consumption and increase the control accuracy. The thrusters configuration generally contains information about thrusters fixed on the satellite body structure, including their location, orientation. One important factor playing a key role in thrusters’ configuration design is satellite force-torque analysis. The proposed configuration, however, should lead to fulfill specified attitude maneuver when the set of force and torque produced by satellite thruster system is adequate. For this purpose, two optimization methods using genetic algorithm (GA) and differential evolution (DE) has been applied to determine the optimal thrusters configuration on the communication satellite body. The cost function employed to minimize both the fuel consumption and error generated by thrusters installation and uncertainties. Moreover, this work allows applying some different constraints in the proposed formulation including minimization of the thruster plume impingement effect on the satellite outer structure and on the solar arrays and the second one is the satellite dimension and geometry. Simulation results show that DE outperforms GA in terms of accuracy and CPU time. Effectiveness of differential evolution algorithm is illustrated in the paper when compared with GA results.

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In this paper, LQG/LTR controller is designed for attitude control of the geostationary satellite at nominal mode. Usage actuator in this paper is the reaction wheel and control torque is determined by the LQR regulator. Usage sensors in this article are sun and earth sensors and EKF are used for estimation of noisy states. LQR controller signal has good performance, if all system's states are considered in system output feedback. But this method is ideal and does not include model noise and sensors noise. Therefore, LQG and LQG/LTR controllers are designed based on the estimated states, and are compared with LQR controller. Controllers gain coefficients are obtained based on linearization about working point. It caused to robustness and similarity of LQG and LQG/LTR response. The results show that control overshoot of LQR is greater than the others.

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This paper is focused upon the development of an efficient multi limit-state reliability analysis method based on extended cross-entropy. In order to achieve a reliable state in the modern critical structural designs, it is necessary to utilize an efficient reliability analysis method to estimate the probability of failure. Reliability analysis of such designs is faced with several challenges, such as rarity of the probability of failure and multiplicity of limit-state functions. To address these issues, in this research an importance sampling (IS)-based method is developed which uses cross-entropy (CE) method to reduce sampling variance, and leads to a fast and accurate reliability analysis method. The main advantage of this method is the ability to perform reliability estimation (especially for rare events) without the most probable failure point (MPP) calculation. The proposed method is demonstrated on several test problems and the results are compared together. Results obtained show that method introduced in this paper provides an effective way of improving reliability analysis of multi limit-state functions. In addition, using the proposed method, prevents the mistake of over-designing while retains usability of the method.

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In this paper, multi-objective Genetic Algorithm has been employed to find the communications satellite optimal transfer trajectory from geosynchronous transfer orbit (GTO) to geosynchronous orbit (GEO) as the destination operational zone. Because of satellite high specific impulse, Orbital maneuvers are considered as impulsive maneuvers in this paper and its validation is verified by comparison of results with continues one. Number of intermediate orbits has been found in order to physical constraints. In this systematic approach, both minimum-time and fuel-saving using constant acceleration simultaneously as the problem strategy to find the optimal transfer trajectory between two orbits. Set of optimal trajectories are plotted in Pareto Front and transfer trajectory can be selected from these points. Finally, the results are compared with the results obtained using STK software and good agreement has been revealed. The satellite access to ground station (GS) is analyzed and optimal position for GS is found by defining lack of access and mass consumption as objective functions. Finally, disturbance torques induced by liquid apogee motor (LAM) is investigated and thrusters' fuel mass for this disturbance torque control is proposed.

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In this paper, a LQG/LTR controller is proposed for attitude control a geostationary satellite at nominal phase. Basically, proposed methodology includes three parts: LQR regulator, EKF, and loop transfer recovery. Controller design is based on the linearized equations of the spacecraft dynamics using reduced quaternion model. Reduced quaternion model solve uncontrollable problem in some subspaces in the linearized state space quaternion model using all four components of quaternion. Spacecraft actuators are reaction wheels and attitude determination sensors are sun and earth sensors. LQR controller is ideal and it doesn’t account for the model uncertainty and sensor noise and it uses the feedback of the full states. To consider the model uncertainty and sensor noise, we have designed EKF which is used by LQG and LQG/LTR controllers. Controller gain coefficients are obtained using a reduced quaternion model, and based on linearization around the equilibrium point and the natural frequency of the closed loop system. To increase the robustness of the design with respect to solar radiation disturbance, singular values of LQG are approximated to Kalman filter, in LTR section. The results demonstrate that LQG/LTR performance is better than LQG’s and LQG/LTR has a good robust stability margin with respect to disturbances.

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In this paper an efficient algorithm of system design with reliability approach is presented. Using this algorithm in subsystem designing of special systems may lead to systematically optimized design with reliability parameters attitude. By applying item tree parameters including general subsystem parameters into the algorithm and by knowing failure modes and the occurrence rate of each failure mode, the design may be improved and necessary reconsiderations can be applied in order to prevent or reduce the probability of failure mode. First some basic conceptual definitions including reliability, availability, capacity and failure rate is explained then various reliability analysis methods like Fault Tree Analysis (FTA), Failure Mode Effect Analysis (FMEA), Reliability Block Diagram (RBD) and Markov analysis is discussed. Reliability and availability distribution over different phases of unlocking and deploying mechanisms are illustrated using Windchill Solution. Subsequently the effect of the different ranges of failure rates of added components on reliability, availability and capacity of whole system is investigated. By analyzing the reliability and availability of system for different phases, it was found that the whole system is under stable situation at the end of each phase. Also results showed that the reliability, availability and capacity of whole system increased and reached a stable level by minimizing the failure rate of the redundant components.

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Fracture phenomenon in orthotropic materials, generally associates with fracture process zone (damaged zone) in crack tip vicinity. Determination of Mechanical properties in this region can help to predict the value or even the direction of crack growth in orthotropic materials. This area contains a multitude of micro cracks which cause difficulties in analytical process of the region Also cause energy waste in damaged zone that can affect the material fracture properties. So far, several models have been proposed to determine the mechanical properties of this region, but due to the immense complexity of this region, the results have not been expressed the behavior of this region properly. Moreover, the existence methods have not been verified with new experimental and numerical data, yet. In present research, a new approach based on experimental and numerical results proposed to investigate the orthotropic damaged zone properties. This model, unlike previous models by offering a range for effective elasticity modulus, can determine the mechanical properties of this region for the presence or absence of micro-cracks interaction among them. The proposed model also, validated and compared with experimental and numerical results.

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This research deals with configuration design and layout optimization of communication satellite. First, an approach is proposed to design the configuration of GEO satellite. Since propulsion subsystem in GEO satellite is the massive item, it has a significant impact on satellite configuration. Consequently, it is necessary to consider the propulsion subsystem influence on satellite configuration. Then layout design process of the satellite components which is one of the complex problems in engineering is performed. In this paper, in order to optimize the layout design of satellite components, the algorithm which consists of two stages, primary and detail layout, is proposed. In order to express geometric constraints mathematically, the Finite Circle Method (FCM) is used. For The mathematical expression of performance constraints, the distance constrains related to distance relationships between components have been developed. The hybrid optimization method is proposed to optimize layout design which is a combination of Simulated Annealing optimization and Quasi Newton methods. The optimization method validation is applied on simple test problem. Finally, the proposed algorithm for configuration and optimal layout design is implemented on communication satellite. The results show that product of inertial (objective function) are minimized and considered constrains of communication satellite are satisfied.

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A novel approach is presented for the reconfiguration of satellite constellations based on Lambert’s theorem. The reconfiguration problem, in this article, is considered with the constraint of overall fuel cost minimization. Hence, orbital maneuvers required for the operation of reconfiguration is designed in such a way that, transferring the satellites to the desired configuration of constellation will be possible at minimal cost. Also, the introduced method of orbital transfer for implementing the reconfiguration phase of satellite constellation has no limitation on the shape and orientation of initial and target orbits such as: co-planarity, coaxiality, circularity and/or the existence of a common point. Moreover, a method is offered for modeling the cost function of reconfiguration problem in which the two important tasks of optimal orbital transfer of satellites to the target configuration of constellation and optimal assignment of each satellite to a specific terminal position or final orbit will be done in one single step. For this purpose and in order to achieving the globally optimal solution of the reconfiguration problem of constellation the hybrid PSO/GA is used. Finally, two different scenarios of reconfiguration of satellite constellation will be modeled once by the presented approach and once by considering determined positions of flight and deployment for the satellites. The obtained results indicate the superiority of the idea presented in this article.

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In the present study, a new attitude stabilization concept has been investigated for a satellite considering failure in one or more reaction wheels. In this approach control torques could be generated using only one thruster mounted on a two axis gimbal mechanism. In the other word, in the absence of reaction wheel(s), control torques are generated by applying a thruster rotating mechanism which can be turned around two axes by thruster vector. If any failure happened in reaction wheels, gimbal angles mechanisms will be added to the system as input controlling. Controller algorithm based on dynamic and kinematic equations of the satellite’s motion, has been developed in the presence of disturbances. Three-axis stabilization of the attitude in a LEO orbit satellites under disturbances has been executed by applying three reaction wheel actuators to produce torque in each direction. Disturbance torques that are commonly applied to the satellites are gravity gradient, solar radiation pressure and aerodynamics. For training the intelligent neuro-fuzzy controller, PID controller is employed. Numerical simulations show that, the recommend controlled method have acceptable results (in the presence of disturbances) and adding of a thruster actuator to the system as a redundancy, could enhance the space missions reliability and if any fault happened in the operation of reaction wheels, thruster mechanisms come in to control system , accurately, and sustained satellite stability at desirability attitude.

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For formation flying of two satellites in a satellite constellation, the relative motion and attitude determination algorithms are the key components that affect the quality of flight and mission efficiency. In this paper orbital relative motion of two satellites with arbitrary Keplerian elliptic orbit and in large distance will be analyzed and also exact and efficient solution for relative motion of the satellite with j2 perturbation which is one of the important perturbation in Low Earth Orbit (LEO) using spherical geometry is proposed. Direct geometric method using spherical coordinates are utilized to achieve this solution. In this method relative position and relative velocity of two satellites are calculated in the satellite constellation based on orbital elements. The obtained results from simulation with STK software, comparison of results with extracted results from equations for satellite with different eccentricity and analysis of the proposed method’s accuracy and fault show that the solution obtained from the geometric method presents the relative motion of the satellite with high accuracy. Thus, the proposed solution will be applicable and effective for relative motion of constellation satellites in space missions.

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Coverage of ground stations by satellites is a very important factor to access geographic, geotechnical and strategic information. This is generally achieved by one or more satellite with specified position and navigation. In this regard, in the area of low altitude orbits regional or global coverage of the Earth's surface is achieved utilizing various mathematical methods to change the position and arrangement of satellites. In this study, the arrangement of certain number of satellites is performed to reach maximum coverage. It is assumed that the satellite constellation is in the symmetrical Walker pattern. In this regard, taking into account the situation of user and determining the initial position of satellite in system, Geometric Dilution of Precision (GDOP) parameters are calculated utilizing a new model. The innovation in this new presented model is employing GDOP in an inverse manner. GDOP is a geometric standard that the less related values for it represents more accuracy in determining the amount of coverage. In this study the effects of compression of the earth as well as chamfer are considered. The calculations are presented for specific geographic areas and only for one day. The results show that by taking advantage of new computing model, the coverage area will dramatically increase. By organized employing of all the satellites in the constellation, with the best received Information from satellites, better coverage can be achieved.

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The purpose of this research is design of solar panels for a satellite which put in to geostationary orbit considering siutable reliability. The process of solar panel design is conducted according to the Design Structure Matrix (DSM) method. In this regard, an initial plan, a subsequent design process improvement, and a final optimized design process are provided. The first level of designed mechanism product tree includes released mechanism, development mechanism, lock and rotation components. Given the importance of ensuring the proper operation of mechanisms in space and reported mission failures due to lack of mechanism’s operation, the reliability network of designed mechanism is constructed and the reliability of designed panel is calculated. The amount of achieved relaiability is then verified according to the mission and system engineering requirements. Nessecary changes are applied on initial design to achieve into the satisfactory reliability for whole solar mechanism. In this regard, the critical paths in reliability network which lead to reduced reliability are investigated, and improvement of the critical path are proposed, to the extend of increasing reliability by discarding redundant components for critical parts.

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In this research, a general mixed mode I/II fracture criterion is developed for fracture investigation of orthotropic materials. Various experimental tests show that cracks always propagate in an isotropic medium and along fiber direction in orthotropic materials. With a novel material model titled an Equivalent Reinforced Isotropic Model (ERIM), fracture criterion can be extended for investigation of fracture in orthotropic materials. This inspires that fracture in orthotropic materials follows the fracture mechanism in isotropic materials. This new criterion is developed based on extension of MTS which is widely used for isotropic materials. Also in this research the effects of T-stress in fracture of some specimens has been studied. A comparison between available experimental observations and theoretical estimation implies on capability of developed criterion for predicting both crack propagation direction and fracture instance, wherein the achieved fracture limit curves are also compatible with fracture mechanism of orthotic materials. It is also shown that non-singular T-stress term has a significant impact on orthotropic material failure, especially when the second mode is dominant mode. It is shown that unlike isotropic materials, fracture toughness of orthotic materials in mode I (K_IC) cannot be introduced as the maximum load bearing capacity and thus new fracture mechanics property, named here as maximum orthotropic fracture toughness in mode I (├ K_IC ┤|ortho) is defined. Considering ease of access, wood is used as experimental specimen for the purpose of comparing the results.

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In hovering, the deputy satellite must use fuel regularly to maintain a constant distance from a source point. Since the amount of fuel in satellites is a key element, it has also a great value to optimize its consumption. The system speed and the time of reaching stability are also other important elements that have been studied in this study. For this purpose, three useful controllers are used. These controllers are LQR controller, Pole Placement and sliding mode controller. In the following, these controllers were optimized using particle swarm optimization algorithm in order to compromise the amount of fuel consumption and the time of reaching stability. Comparing the final results shows that the sliding mode controller can be the best option for optimizing hovering system.

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The satellites on the ground during construction and transportation, in launching stage and operation in space are under various types of dynamic loads, including high and low frequency vibrational loads, acoustics, shock, impact, etc., each of which can be an important source in the creation of stress on the satellite. The satellite components should be designed in such a way that can continue to operate while facing these situations. Electronic boards, in particular their solder joints, are critical components of satellites. Therefore, investigation of damage in design process of boards have great importance. Loading pattern on the satellite during its operation is usually random which considered as quasi-static load. Improvement of the design of the satellite against the weaknesses shown while facing different loads is essential, and given the fact that it is time consuming and costly to carry out laboratory tests, the use of analytical methods for checking the strength and lifetime of the structure can be very useful. In this research, random vibrations environment is equivalent to pseudo-static loads, and using the multilayer plate theory, the stresses in solder joints and failure of joints under this loading will be investigated. Also, the effect of parameters such as electronic board width and the boundary condition of the printed circuit board on the solder joints' stress will be considered in analytical solution.

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The modern business environment is constantly changing, and manage this change to adapt to a future of uncertainties is a challenge that necessitates supply chain flexibility. We introduce a method for measuring flexibility based on gray system. For this purpose, the flexibility of Alpha Automobile Company that operates in Iran's automobile industry has been measured. The results suggested that sourcing limitations and Distribution problems are the most serious vulnerabilities that threaten Alpha Automobile Company. Therefore, the company must plan to improve its flexibility by selecting the appropriate set of capabilities. Moreover, effectiveness , supply flexibility and flexibility in order fulfillment were identified as three important capabilities of Alpha Automobile Company. Keywords: Flexibility; Gray system; Vulnerabilities; Capabilities The modern business environment is constantly changing, and manage this change to adapt to a future of uncertainties is a challenge that necessitates supply chain flexibility. We introduce a method for measuring flexibility based on gray system. For this purpose, the flexibility of Alpha Automobile Company that operates in Iran's automobile industry has been measured. The results suggested that sourcing limitations and Distribution problems are the most serious vulnerabilities that threaten Alpha Automobile Company. Therefore, the company must plan to improve its flexibility by selecting the appropriate set of capabilities. Moreover, effectiveness , supply flexibility and flexibility in order fulfillment were identified as three important capabilities of Alpha Automobile Company. Keywords: Flexibility; Gray system; Vulnerabilities; Capabilities